High-power liquid-propellant rocket engines are known in rocket engineering and they are widely used as parts of launchers designed for various purposes.
Known in the art under code name RD-253 is a liquid-propellant rocket engine (LRE) designed by the "Energomash" Scientific and Production Corporation (NPO Energomash) (see in the encyclopaedia: "Cosmonautics", Editor-in-Chief V. P. Glushko, Moscow, 1985, pp. 330-331). This LRE comprises a combustion chamber, a gas generator, a turbopump assembly, automatic equipment, pipelines that interconnect hydraulically engine assemblies with each other, a gas line that connects the turbine of the turbopump assembly to the combustion chamber, units for fastening the LRE to a launcher so as to make it capable of turning (swinging) in a vertical plane.
A disadvantage of this prior art LRE design consists in that it can ensure turning (swinging) of the LRD, as required for changing its thrust vector direction, only in a single plane. It should be also pointed out that large size of the bellows balances used in this prior art design leads to an increased axial overall dimension of the engine.
These disadvantages of the prior art design also exist in the design of a dual-chamber LRE which also has a considerable axial overall dimension. The above disadvantages prevent the production of a launcher having smaller mass and overall dimensions.
Also known in the art is an F-1 LRE designed by Rocketdyne Co. of the U.S.A. (see in the book: TsIAM, "Foreign Aircraft and Rocket Engines", 1971, pp. 436-439). This LRE comprises a combustion chamber, a gas generator, a turbopump assembly, automatic equipment, inner pipelines of the engine, and a gimbal assembly.
As applied to a dual-chamber engine design, this prior art design requires the development of a special frame. In a number of cases, it is also difficult to ensure smaller axial overall dimensions for this engine and, hence, to optimize the launcher as far as its overall dimensions and mass are concerned.
The closest to the liquid-propellant rocket engine of the present invention is an LRE under code name RD-219 designed by NPO Energomash (see in the encyclopaedia: "Cosmonautics", Editor-in-Chief V. P. Glushko, Moscow, 1985, p. 330). This prior art liquid-propellant rocket engine comprises two combustion chambers fixed to a frame, a turbopump assembly fastened also to the frame and having a turbine, oxidizer and fuel pumps, and pipes for feeding the oxidizer and fuel to a gas generator and to the combustion chamber of the engine.
A limitation of this prior art design consists in that the engine chambers are rigidly fixed to the frame. They cannot turn so as to change the thrust vector direction and, in a number of cases, they require to provide special control spaces on board the launcher. Besides, this engine has no pipe between the turbine and the combustion chambers so that gas is exhausted from the gas generator overboard, instead of contributing to an increase in the specific thrust developed by the rocket engine as the case is for a system with gas from the gas generator being afterburned. If this engine is made with "swinging" combustion chambers using the system for afterburning the gas, then both the axial and diametric overall dimensions of the design will become larger.